This disclosure relates to gas turbine engines and particularly to internally cooled rotor blades.
A gas turbine engine typically includes a fan section, a compressor section, a combustor section and a turbine section. Air entering the compressor section is compressed and delivered into the combustor section where it is mixed with fuel and ignited to generate a high-speed exhaust gas flow. The high-speed exhaust gas flow expands through the turbine section to drive the compressor and the fan section. The compressor section typically includes low and high pressure compressors, and the turbine section includes low and high pressure turbines.
As is well known, the aircraft engine industry is experiencing a significant effort to improve the gas turbine engine's performance while simultaneously decreasing its weight. The ultimate goal has been to attain the optimum thrust-to-weight ratio. One of the primary areas of focus to achieve this goal is the “hot section” of the engine since it is well known that engine's thrust/weight ratio is significantly improved by increasing the temperature of the turbine gases. However, turbine gas temperature is limited by the metal temperature constraints of the engine's components. Significant effort has been made to achieve higher turbine operating temperatures by incorporating technological advances in the internal cooling of the turbine blades.
Serpentine core cooling passages have been used to cool turbine blades. The serpentine cooling passage is arranged between the leading and trailing edge core cooling passages in a chord-wise direction. One typical serpentine configuration provides “up” passages arranged near the leading and trailing edges fluidly joined by a “down” passage. This type of cooling configuration may have inadequacies in some applications. To this end, a double wall cooling configuration has been used to improve turbine blade cooling.
In a double wall blade configuration, thin hybrid skin core cavity passages extend radially and are provided in a thickness direction between the core cooling passages and each of the pressure and suction side exterior airfoil surfaces. Double wall cooling has been used as a technology to improve the cooling effectiveness of a turbine blades, vanes, blade out air seals, combustor panels, or any other hot section component. Often, core support features are used to resupply air from a main body core, which creates the core passages, into the hybrid skin core cavity passages, which creates the skin passages.
With traditional double wall configurations, a cooling benefit is derived from passing coolant air from the internal radial flow and/or serpentine cavities through the “cold” wall via impingement (resupply) holes and impinging the flow on the “hot” wall. These core support (resupply) features are typically oriented perpendicular to the direction of flow in the hybrid skin core cooling cavities. These perpendicular core support (resupply) features induce local flow vortices which generate a significant amount of turbulent mixing to occur locally within the hybrid skin core cavity passage. Although the impingement flow field characteristics associated with the resupply holes may appear beneficial, they create local flow characteristics which are not advantageous from an internal cooling perspective. Adverse impacts due to disruptive impingement resupply features oriented perpendicular to the downstream streamwise flow direction with in the hybrid skin core cavity generate pressure and momentum mixing losses that mitigate the favorable convective cooling flow field characteristics. Potential improvements in the internal flow field cooling qualities are diminished due to the disruptive nature of the injection of high pressure and velocity resupply cooling air flow normal to main hybrid skin core cooling passage flow direction. The potential decrease in bulk fluid cooling temperature may be adversely impacted by the additional cooling air heat pickup incurred due to the high impingement heat transfer and subsequent heat removal from the exterior hot wall. In a purely convective hybrid skin core cooling channel passage the locally high impingement heat transfer generated by the resupply features oriented normal to the downstream cooling flow may produce large local metal temperature gradients that result in locally high thermal strain and subsequent thermal mechanical fatigue crack initiation and propagation failure mechanisms.
Improving the mixing characteristics of the two different flows through the incorporation of “in-line” or “angled” resupply orientation and unique geometric features can improve the overall convective cooling characteristics of the internal flow field and increase the thermal cooling effectiveness of resupplied hybrid skin core cooling cavity passages. The intent of this invention is improve the relative alignment of the resupply cooling flow with the downstream cooling flow within the hybrid skin core cooling channel passages. Additionally it is also desirable to introduce the resupply cooling flow at a mass and momentum flux ratio that is ≥ the mass and momentum flux of the downstream cooling flow within the hybrid skin core cooling channel passage immediately adjacent to the internal surface of the hot exterior airfoil wall. By introducing resupply flow through a diffused geometric feature the relative mass and momentum mixing of the two different flow streams is more easily controlled by modifying the expansion ratio and geometry shape of the diffusing section of the resupply geometry.